Patents U.S. Pat. No. 5,498,129 and FR 2 719 553 describe such a structure that is secured to the rear end of a tail boom extending the fuselage and that receives a rotor for opposing the torque exerted on the fuselage by the main rotor for providing the aircraft with lift and propulsion.
That structure comprises an annular (or tubular) wall forming a duct defining a flow tunnel for the stream of air driven by the anti-torque rotor which is housed in the duct.
In general, such a duct presents in succession, from upstream to downstream: a converging inlet portion; a cylindrical portion (constant radius); and a diverging portion that terminates by a tunnel outlet having a rounded edge; each end of the duct includes an annular flange or collar.
The tail structure also includes two side flanks (a left flank and a right flank) or fairing side walls, each having an orifice and secured via the edge of said orifice to a respective one of the two flanges of the tubular duct.
The tail structure further includes a substantially vertical fin or stabilizer secured to the side walls of the fairing and extending above them, substantially along a fore-and-aft vertical plane of general symmetry of the aircraft fuselage, or else inclined at an angle relative to said plane.
Such a tail structure is generally obtained by assembling at least four main parts, to which there need to be added reinforcing parts extending between the side walls of the fairing that they unite, and also at least one connection part for securing said tail structure to the tail boom.
Those various parts are generally made of a composite material comprising an organic matrix and reinforcing fibers; a portion of the fairing side walls and of the annular duct wall described in those patents further comprises a honeycomb layer (or the equivalent).
In order to make each of those parts, the usual technique is to place one or more layers of a preimpregnated fabric or “prepreg” on a male mold (projecting in relief) or in a female mold (recessed) of a shape that matches that of the part, the prepreg generally being constituted essentially by fibers (e.g. carbon fibers) coated with a thermosetting resin; during that “draping” operation, the fibers are oriented along one or more predetermined directions in order to obtain the desired mechanical characteristics for the part.
When the part that is to be made includes a thick core such as a honeycomb, the procedure generally comprises a first draping operation to form a skin (internal or external), followed by depositing the core, and then generally a second draping operation covering the core for the purpose of forming a second skin (respectively external or internal).
The preform as obtained in that way is then made rigid by applying heat, and where appropriate a vacuum, in an autoclave or oven in which a complete mold (male and female) is placed that surrounds the preform; that step serves to polymerize or cross-link the organic matrix.
The operations of making the preform are generally performed manually, thereby leading to losses of composite material, and it does not enable the mechanical characteristics of the parts in a series of parts of identical shape to be made in completely reproducible manner.
In order to mechanize the draping operation, it is known to use a fiber placement technique whereby the preimpregnated fibers are deposited by a machine on the outside surface of a mandrel mounted to rotate about an axis of rotation, as described in particular in patents FR 2 766 407 and U.S. Pat. No. 6,613,258.
Nevertheless, that technique is generally restricted to making parts of convex shape and presenting circular symmetry, which does not apply to a tail structure for a rotary wing aircraft with a ducted tail rotor. In particular, the front zone where the tail and the body of the fairing join together presents shapes that are complex, having multiple curvatures: zones in which curvature is reversed where a concave wall is adjacent to a convex wall, and zones of small radius of curvature.
The manufacture of such a tail structure then requires the various composite material parts to be assembled together, by applying adhesive to portions of these parts that are placed in mutually overlapping positions, by riveting, by adhesive and riveting, and/or by using other bonding means.
Those bonding techniques likewise are not suitable for obtaining a resulting assembled structure presenting mechanical characteristics that are completely reproducible. Those bonding techniques also lead to a harmful increase in the weight of the resulting assembled structure.